The present invention generally relates to turbomachinery, and more particularly to methods and hardware capable of cooling turbomachinery components while reducing thermally-induced strains and stresses that result from components of an assembly exhibiting different amounts of thermally expansion.
Higher operating temperatures for gas turbines, including aircraft gas turbine engines and land-based gas turbine engines used in the power-generating industry, are continuously sought in order to increase their efficiency. However, as operating temperatures increase, the high temperature durability of the engine components must correspondingly increase. Significant advances in high temperature capabilities have been achieved through the formulation of iron-, nickel- and cobalt-based superalloys. Nonetheless, superalloy components must often be air-cooled and/or protected by some form of thermal and/or environmental coating system in order to exhibit suitable service lives in certain sections of a gas turbine engine, such as the turbine, combustor and augmentor.
As an example, FIG. 1 represents an axial cross-section of a nozzle segment 10 of a land-based gas turbine engine. The nozzle segment 10 is one of a number of nozzle segments that are assembled together to form an annular-shaped high pressure turbine (HPT) nozzle assembly of the turbine engine. The segment 10 is made up of at least one vane 12, which defines an airfoil and extends between inner and outer sidewalls 14 and 16 (also referred to as platforms or bands). The vane 12 and sidewalls 14 and 16 can be formed separately and then assembled, such as by brazing the ends of the vane 12 within openings defined in the sidewalls 14 and 16. Alternatively, the entire segment 10 can be formed as an integral casting. The vane 12 and sidewalls 14 and 16 can be formed of such conventional materials as nickel-, cobalt-, or iron-based superalloys of types suitable for use in gas turbine engines.
As a result of being located in the high pressure turbine section of the engine, the vane 12 and the surfaces of the sidewalls 14 and 16 facing the vane 12 are subjected to the hot combustion gases from the engine's combustor. A thermal barrier coating (TBC) system may be applied to the surfaces of the vane 12 and sidewalls 14 and 16 exposed to the hot combustion gases to provide environmental protection and reduce heat transfer to the segment 10. Alternatively or in addition, compressor bleed air may be supplied to the vane 12 and sidewalls 14 and 16 to provide forced air cooling, such as by an impingement or film cooling technique. As an example, FIG. 1 represents the use of an impingement cooling technique on both the inner and outer sidewalls 14 and 16 of the nozzle segment 10, and FIG. 2 provides a more detailed view illustrating impingement cooling of the sidewall 16. In FIG. 1, impingement plates 18 and 20 are shown as coupled to the sidewalls 14 and 16 to create a cavity or chamber 22 therebetween. Bleed air is drawn from the engine's compressor (not shown) and supplied to the sides of the impingement plates 18 and 20 facing away from their respective sidewalls 14 and 16. Numerous small apertures, sometimes referred to as impingement cooling holes 24 (FIG. 2), are present in the plates 18 and 20 that direct the bleed air in a normal direction toward the surfaces of the sidewalls 14 and 16 opposite the vane 12, achieving what is referred to as backside cooling of the sidewalls 14 and 16. FIG. 2 further shows the outer sidewall 16 as having film cooling holes 26 through which the cooling air within the chamber 22 is discharged at an acute angle to the surface of the sidewall 16 facing the hot gas path of the engine to achieve a film cooling effect at that surface. Impingement and film cooling techniques are well known in the art, and therefore do not require further explanation.
As one would expect, the nozzle segment 10 expands and contracts when heated and cooled, respectively, during transient engine operating conditions. Because they are in direct contact with the hot combustion gases, the vane 12 and sidewalls 14 and 16 sustain temperatures that are significantly higher than the mounting hardware to which the sidewalls 14 and 16 are attached. Because both surfaces of the impingement plates 18 and 20 are directly contacted by the cooling air, the plates 18 and 20 also tend to be at lower temperatures than the sidewalls 14 and 16. As a result, the sidewalls 14 and 16 typically expand and contract more than the impingement plates 18 and 20.
The impingement plates 18 and 20 are often fabricated from thin sheet metal to minimize their weight and simplify the creation of their cooling holes 24. The plates 18 and 20 are typically attached to their respective sidewalls 14 and 16, often with welds or some combination of welds and clamps. As a nonlimiting example, FIG. 3 schematically represents a plan view of the impingement plate 20 nested within a recess 28 defined in the outer sidewall 16 and surrounding the chamber 22 (not seen), and represents the periphery 30 of the plate 20 being attached with spot welds 32 to the sidewall 16. The inner sidewall 14 and its impingement plate 18 can be similarly configured to that shown in FIG. 3 for the outer sidewall 16 and plate 20. FIG. 4 is a partial cross-section of FIG. 3 representing one of the welds 32. The periphery 30 of the plate 20 is represented as lying in a separate plane from the bulk of the plate 20, with roughly an S-shaped wall 34 therebetween.
As the sidewall 16 thermally expands, the sidewall 16 will move leftward in FIG. 4 a greater distance than the periphery 30 of the plate 20 due to differences in the temperatures of the sidewall 16 and plate 20, as explained above. Because the weld 32 rigidly attaches the plate 20 to the sidewall 16, strains and stresses are induced in the weld 32 as well as the plate 20. Consequently, the plate 20 and weld 32 are both prone to damage from thermally-induced strains and stresses that occur as the sidewall 16 thermally expands during high temperature excursions. As reported in U.S. Pat. No. 4,693,667, which discloses an impingement plate having a similar S-shaped portion along its perimeter, the S-shaped wall 34 of the plate 20 is capable of accommodating differences in thermal expansion between the plate 20 and sidewall 16 to some degree. However, experience has shown that relatively thin impingement plates 18 and 20 and welds 32 of the types represented in FIGS. 3 and 4 are nonetheless susceptible to cracking and fragmenting.
Because cracks and voids in the plate 20 provide additional passages through which bleed air can flow through the plate 20, damage to the plate 20, though not likely to be pose a direct structural concern, can lead to oxidation and cracking of the sidewall 16 due to a lack of efficient cooling air flow toward the sidewall 16. In addition, cracks and voids in the plate 20 can result in excessive loss of bleed air and reduce the overall efficiency of the engine. Therefore, there is a need for more robust designs for impingement plates to reduce their likelihood of cracking during engine operation.